The present invention relates generally to gas turbine engines and, more specifically, to turbines therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through multiple turbine stages. A turbine stage includes a stationary turbine nozzle having stator vanes which guide the combustion gases through a downstream row of turbine rotor blades extending radially outwardly from a supporting disk which is powered by extracting energy from the gases.
A first stage or high pressure turbine nozzle first receives the hottest combustion gases from the combustor which are directed to the first stage rotor blades which extract energy therefrom. A second stage turbine nozzle is disposed immediately downstream from the first stage blades and is followed in turn by a row of second stage turbine rotor blades which extract additional energy from the combustion gases.
As energy is extracted from the combustion gases, the temperature thereof is correspondingly reduced. However, since the gas temperature is relatively high, the high pressure turbine stages are typically cooled by channeling through the hollow vane and blade airfoils cooling air bled from the compressor and/or are made of high temperature capability and high heat resistant materials. The greater the temperature capability of the turbine materials are, the more expensive the turbine airfoils are. Since the cooling air is diverted from the combustor, the overall efficiency of the engine is correspondingly reduced. It is therefore highly desirable to minimize the use of such cooling air for maximizing overall efficiency of the engine and reduce the expense of the airfoils by using turbine materials having lower heat resistance properties.
The amount of cooling air required is dependent on the temperature of the combustion gases. That temperature varies from idle operation of the engine to maximum power operation thereof. Since combustion gas temperature directly affects the maximum stress experienced in the vanes and blades, the cooling air requirement for the turbine stages must be effective for withstanding the maximum combustion gas temperature operation of the engine although that running condition occurs for a relatively short time during engine operation.
For example, a commercial aircraft gas turbine engine which powers an aircraft in flight for carrying passengers or cargo experiences its hottest running condition during aircraft takeoff. For a military aircraft engine application, the hottest running condition depends on the military mission, but typically occurs during takeoff with operation of an afterburner. And, for a land-based gas turbine engine which powers an electrical generator, the hottest running condition typically occurs during the hot day peak power condition.
The maximum combustion gas temperature therefore varies temporally over the operating or running condition of the engine. The maximum combustion gas temperature also varies spatially both circumferentially and radially as the gases are discharged from the outlet annulus of the combustor. This spatial temperature variation is typically represented by combustor pattern and profile factors which are conventionally known. The highest temperature environment occurs in a portion of the gas flow where hot streaks from the combustor causes temperature variations in upstream airfoil rows in the stator vanes or nozzles and in the rotating blades. Unsteadiness caused by upstream wakes and hot streaks causes a pattern in the gas flow through the turbine where the airfoil is hot, the pressure side is hotter than the suction side, and the middle of the passage is cold. In a rotor, a variation circumferentially in the absolute frame causes an upstream unsteady disturbance. Cold nozzle wakes and combustor hot streaks produce large unsteady temperature variations.
Accordingly, each turbine stage, either blades or vanes, is typically specifically designed for withstanding the maximum combustion gas temperature experienced both temporarily and spatially in the combustion gases disposed directly upstream therefrom. Since the airfoils in each row of vanes and blades are identical to each other, the cooling configurations therefor and materials and material properties thereof are also identical and are effective for providing suitable cooling and heat resistance at the maximum combustion gas temperatures experienced by the individual stages for maintaining the maximum airfoil stress, including thermal stress, within acceptable limits for ensuring a suitable useful life of the turbine stages.
It is therefore highly desirable to have a gas turbine engine and its turbine airfoils with reduced temperature capability and where required with reduced cooling requirements.